WO2015042262A1 - Passage de refroidissement sinueux pour composant de moteur - Google Patents

Passage de refroidissement sinueux pour composant de moteur Download PDF

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Publication number
WO2015042262A1
WO2015042262A1 PCT/US2014/056302 US2014056302W WO2015042262A1 WO 2015042262 A1 WO2015042262 A1 WO 2015042262A1 US 2014056302 W US2014056302 W US 2014056302W WO 2015042262 A1 WO2015042262 A1 WO 2015042262A1
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WO
WIPO (PCT)
Prior art keywords
passageway
recited
boas
inlet
gas turbine
Prior art date
Application number
PCT/US2014/056302
Other languages
English (en)
Inventor
Dmitriy A. ROMANV
Original Assignee
United Technologies Corporation
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corporation filed Critical United Technologies Corporation
Priority to US15/023,309 priority Critical patent/US10196931B2/en
Priority to EP14845824.3A priority patent/EP3047113B1/fr
Publication of WO2015042262A1 publication Critical patent/WO2015042262A1/fr

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/20—Actively adjusting tip-clearance
    • F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • F05D2230/211—Manufacture essentially without removing material by casting by precision casting, e.g. microfusing or investment casting
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/15—Two-dimensional spiral
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/25—Three-dimensional helical
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/204—Heat transfer, e.g. cooling by the use of microcircuits
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/209—Heat transfer, e.g. cooling using vortex tubes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • Gas turbine engines include blades configured to rotate and extract energy from hot combustion gases that are communicated through the gas turbine engine.
  • An outer casing of the gas turbine engine may support one or more blade outer air seals (BOAS) that provide an outer radial flow path boundary for the hot combustion gases.
  • BOAS may include cooling passageways configured to route a flow of cooling fluid therein.
  • One known BOAS includes parallel cooling passageways extending between circumferential edges thereof.
  • One exemplary embodiment of this disclosure relates to a gas turbine engine including a component having a body.
  • the body includes a tortuous cooling passageway, which provides a flow path extending between an inlet in a first surface of the body and an exit in a second surface of the body.
  • the flow path includes at least one bend between the inlet and the exit.
  • the inlet is provided about an inlet axis.
  • the tortuous cooling passageway is a three-dimensional spiral passageway, and the flow path moves progressively further away from the inlet axis as the flow path extends from the inlet to the exit.
  • the tortuous cooling passageway is a Z-shaped passageway.
  • the Z-shaped passageway includes three sloped portions, and each of the three sloped portions is successively spaced radially further from the first surface.
  • the cooling passageway is an M-shaped passageway.
  • the M-shaped passageway includes a first inlet and a second inlet converging to a single exit.
  • the component includes a plurality of tortuous cooling passageways, and wherein borders of adjacent cooling passageways are aligned relative to one another in at least one direction.
  • the component is a blade outer air seal (BOAS), and wherein the second surface is positioned adjacent a tip of a rotor blade.
  • BOAS blade outer air seal
  • the BOAS includes a body including a tortuous cooling passageway, which provides a flow path extending in each of a radial, axial, and circumferential direction.
  • the body includes a first surface and a second surface, the flow path provided between an inlet in the first surface and an exit in a second surface.
  • the BOAS includes at least one of a three-dimensional spiral passageway, a Z-shaped passageway, and an M-shaped passageway.
  • the BOAS includes a plurality of three-dimensional spiral passageways, a plurality of Z-shaped passageways, and a plurality of M-shaped passageways.
  • the plurality of Z-shaped and M-shaped passageways are provided adjacent edges of the BOAS to provide a perimeter, and wherein the plurality of three-dimensional spiral passageways are provided within the perimeter.
  • the tortuous cooling passageway includes at least one of trip strips and pedestals therein.
  • the article includes a first portion providing a negative of an inlet, and a second portion providing a negative of an outlet.
  • the second portion is spaced from the first portion.
  • a third portion provides a negative of a tortuous cooling passageway, with the third portion extending between the first portion and the second portion.
  • the third portion includes a plurality of sloped portions, with each of the plurality of sloped portions successively spaced further away from the first portion.
  • the third portion further includes a plurality of legs extending between adjacent ones of the plurality of sloped portions. Each of the plurality of legs are successively spaced further away from the first portion.
  • the tortuous cooling passageway is one of a three-dimensional spiral passageway, a Z-shaped passageway, and an M-shaped passageway.
  • Figure 1 illustrates a schematic, cross-sectional view of a gas turbine engine.
  • Figure 2 illustrates a cross-section of a portion of a gas turbine engine.
  • FIG 3 illustrates an example blade outer air seal (BOAS).
  • BOAS blade outer air seal
  • Figure 4A is a partial, perspective view of the BOAS of Figure 3, sectioned along line 4-4 in Figure 3, and in particular illustrates a three-dimensional spiral cooling passageway.
  • Figure 4B is a partial view of the BOAS of Figure 3, sectioned along line 4-4 in Figure 3.
  • Figure 5 represents the three-dimensional spiral cooling passageway of Figure 4.
  • Figure 6 illustrates an arrangement of adjacent cooling passageways.
  • Figure 7 represents a Z-shaped cooling passageway.
  • Figure 8 represents an M-shaped cooling passageway.
  • Figure 9 represents an example cooling arrangement.
  • Figure 10A represents a divergent cooling passageway.
  • Figure 10B represents a U-shaped cooling passageway.
  • Figure IOC represents an angled cooling passageway.
  • Figure 11A illustrates a cooling passageway with trip strips.
  • Figure 1 IB illustrates a cooling passageway with pedestals.
  • Figure 12 illustrates a casting article used for forming the three- dimensional spiral cooling passageway in one example.
  • FIG. 1 schematically illustrates an example gas turbine engine 20 that includes a fan section 22, a compressor section 24, a combustor section 26, and a turbine section 28.
  • Alternative engines might include an augmenter section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flow path B while the compressor section 24 draws a core airflow C in along a core flow path where air is compressed and communicated to a combustor section 26.
  • air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and utilized to drive the fan section 22 and the compressor section 24.
  • turbofan gas turbine engine depicts a turbofan gas turbine engine
  • the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
  • the concepts disclosed herein can further be applied outside of gas turbine engines.
  • the example engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
  • the low speed spool 30 generally includes an inner shaft 40 that connects a fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46.
  • the inner shaft 40 drives the fan 42 through a speed change device, such as a geared architecture 48, to drive the fan 42 at a lower speed than the low speed spool 30.
  • the high speed spool 32 includes an outer shaft 50 that interconnects a high pressure (or second) compressor section 52 and a high pressure (or second) turbine section 54.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A.
  • a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
  • the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54.
  • the high pressure turbine 54 includes only a single stage.
  • a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure” compressor or turbine.
  • the example low pressure turbine 46 has a pressure ratio that is greater than about five (5).
  • the pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
  • a mid-turbine frame 58 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46.
  • the mid- turbine frame 58 further supports bearing systems 38 in the turbine section 28 as well as setting airflow entering the low pressure turbine 46.
  • the core airflow C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46.
  • the mid-turbine frame 58 includes vanes 60, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid- turbine frame 58. Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.
  • the disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine.
  • the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10).
  • the example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.
  • the gas turbine engine 20 includes a bypass ratio greater than about ten (10: 1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.
  • the fan section 22 of the engine 20 is designed for a particular flight condition— typically cruise at about 0.8 Mach and about 35,000 feet.
  • 'TSFC' Thrust Specific Fuel Consumption
  • "Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
  • the low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non- limiting embodiment the low fan pressure ratio is less than about 1.45.
  • Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/ (518.7 °R)] 05 .
  • the "Low corrected fan tip speed,” as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.
  • Figure 2 illustrates a portion 62 of a gas turbine engine, such as the gas turbine engine 20 of Figure 1.
  • the portion 62 represents the high pressure turbine 54.
  • other portions of the gas turbine engine 20 could benefit from the teachings of this disclosure, including but not limited to, the fan section 22, the compressor section 24 and the low pressure turbine 46.
  • a rotor disk 66 (only one shown, although multiple disks could be axially disposed within the portion 62) is mounted for rotation about the engine central longitudinal axis A.
  • the portion 62 includes alternating rows of rotating blades 68 (mounted to the rotor disk 66) and static vane assemblies 70.
  • the vane assemblies 70 each includes a plurality of vanes 70A, 70B that are supported within an outer casing 69 of the engine static structure 36 ( Figure 1).
  • Each blade 68 of the rotor disk 66 includes a blade tip 68T at a radially outermost portion of the blade 68.
  • the rotor disk 66 is arranged such that the blade tips 68T are located adjacent a blade outer air seal (BOAS) assembly 72.
  • the BOAS assembly 72 may find beneficial use in many industries including aerospace, industrial, electricity generation, naval propulsion, pumps for gas and oil transmission, aircraft propulsion, vehicle engines and stationery power plants.
  • the BOAS assembly 72 is disposed in an annulus radially between the outer casing 69 and the blade tip 68T.
  • the BOAS assembly 72 generally includes a support structure 74 and a multitude of BOAS segments 76 (only one shown in Figure 2).
  • the individual BOAS segments 76 are each individually referred to as a "BOAS segment” or simply a "BOAS.”
  • the BOAS segments 76 may be arranged to form a full ring hoop assembly that circumferentially surrounds the associated blades 68.
  • the support structure 74 is mounted radially inward from the outer casing 69, and includes forward and aft flanges 78A, 78B that receive forward and aft attachment hooks 76A, 76B of the BOAS segments 76.
  • the forward and aft flanges 78A, 78B may be manufactured of a material such as a steel or nickel-based alloy, and may be circumferentially segmented for the receipt of the BOAS segments 76.
  • a secondary cooling airflow S may be communicated to the BOAS segments 76.
  • the secondary cooling airflow S can be sourced from the high pressure compressor 52 or any other portion of the gas turbine engine 20.
  • the secondary cooling airflow S provides a biasing force that biases the BOAS segment 76 radially inward toward the engine central longitudinal axis A.
  • the forward and aft flanges 78A, 78B are portions of the support structure 74 that limit radially inward movement of the BOAS segment 76 and that maintain the BOAS segment 76 in position.
  • FIG. 3 illustrates a perspective view of an example BOAS segment 80 according to this disclosure. While BOAS segments are discussed herein, it should be understood that this disclosure extends to other engine components, such as blades and vanes, as examples.
  • the BOAS segment 80 includes a fore edge 82, an aft edge 84, and a main body portion 86 extending axially (e.g., relative to the engine central longitudinal axis A, or the "axial direction A") therebetween.
  • the main body portion 86 includes a plurality of cooling passageways receiving a portion of the secondary cooling airflow S, as will be discussed in detail below.
  • the BOAS segment 80 includes attachment hooks 88, 90, 92, 94, 96, 98, 100, which extend upwardly from the main body portion 86 adjacent the aft edge 84.
  • the attachment hooks 88, 90, 92, 94, 96, 98, 100 are shown for illustrative purposes only and are not intended to limit this disclosure.
  • the BOAS segment 80 further includes a first circumferential edge 102, and a second circumferential edge 104.
  • the main body portion 86 includes a radially outer surface 106, and a radially inner surface 108.
  • the radially outer and inner surfaces 106, 108 are spaced- apart from one another in the radial direction Z, which is normal to the engine central longitudinal axis A.
  • the main body portion 86 further includes a plurality of tortuous cooling passageways configured to communicate the secondary cooling air flow S between the radially outer and radially inner surfaces 106, 108.
  • tortuous refers to a cooling passageway that provides a flow path having at least one bend or turn between an inlet and an exit thereof. Several example tortuous cooling passageways are discussed herein.
  • a first example cooling passageway is illustrated at 110.
  • the cooling passageway 110 provides a three-dimensional spiral flow path 112 between the inlet 114 and the exit 116 in the radially inner surface 108.
  • the cooling passageway 110 is arranged such that the flow path 112 moves progressively farther away from an inlet axis 118 as the flow path 112 moves from the inlet 114 to the exit 116.
  • the cooling passageway 110 is in fluid communication with the inlet 114.
  • the inlet 114 is provided about the inlet axis 118, which in this example is arranged parallel to the radial direction Z.
  • the cooling passageway 110 Moving radially inward from the inlet 114, the cooling passageway 110 includes a first sloped portion 120 turning the flow path 112 from a generally radial direction Z to a generally axial direction A, and configured to direct a secondary cooling flow S toward a first leg 122 of the flow path 112.
  • the first leg 122 extends in a circumferential direction Y, which is substantially normal to the axial direction A.
  • the first leg 122 is in communication with a second sloped portion 124, which runs substantially parallel to the first sloped portion 120.
  • the second sloped portion 124 leads to a second leg 126, which extends in the circumferential direction Y and, in turn, leads to a third sloped portion 128.
  • the third sloped portion 128 extends in the axial direction A toward a third leg 130, which extends circumferentially to a fourth sloped portion 132.
  • the fourth sloped portion 132 is in fluid communication with the exit 116.
  • Each sloped portion 120, 124, 128, 132 is inclined (or, angled) to extend non-parallel to the radially outer surface 106 to direct the secondary cooling flow S radially toward the exit 116. That is, in the example of Figures 4A-4B, as the secondary cooling flow S travels along each sloped portion 120, 124, 128, 132, the secondary cooling flow S travels both axially along the length of the particular sloped portion and radially toward the exit 116. Accordingly, each successive sloped portion 120, 124, 128, 132 is radially spaced (e.g., in the direction Z) further from the radially outer surface 106 than the prior sloped portion.
  • each successive leg 122, 126, 130 is radially spaced further from the radially outer surface 106 than the prior leg. It should be understood that the legs 122, 126, 130 may also be sloped (e.g., inclined to extend non-parallel to the radially outer surface 106) alternatively, or in addition to, the sloping of the sloped portions 120, 124, 128, 132. [0068] It should further be understood that while four sloped portions 120, 124, 128, 132 and three legs 122, 126, 130 are illustrated, the cooling passageway 110 could include any number of sloped portions and legs.
  • a portion of a secondary cooling flow S is routed into the cooling passageway 110, and flows along the flow path 112 to cool the BOAS segment 80.
  • the secondary cooling flow S exits the cooling passageway 110 out the exit 116, and generates a film providing additional sealing between the BOAS segment 80 the adjacent blade tips 68T.
  • the exit 116 may be shaped to provide a desired film.
  • Figure 5 schematically represents the cooling passageway 110 viewed from a location radially outboard of the radially outer surface 106.
  • the flow path 112 directs the secondary cooling flow S in three directions, radially (in direction Z) between the outer surface 106, and the inner surface 108, axially, via the sloped portions 120, 124, 128, 132, and circumferentially, by way of the legs 122, 126, 130. This provides a relatively large effective cooling area in a relatively small three- dimensional space.
  • the main body portion 86 of the BOAS may include a plurality of the cooling passageways 110 positioned adjacent one another.
  • circumferential borders of adjacent passageways may be circumferentially aligned. That is, with reference to Figure 6, the circumferential border of the fourth sloped portion 132 of the cooling passageway 110A is spaced a circumferential distance D from a circumferential border of the third sloped portion 128 of the adjacent cooling passageway HOB. In one example, the distance D is zero, in which case the circumferential borders of the cooling passageways 110A, HOB are circumferentially aligned. This relatively close packing between adjacent cooling passageways 110A, HOB is possible due to the third sloped portion 132 being radially spaced from the second sloped portion 128, as described above.
  • axial borders of adjacent cooling passageways may be axially aligned.
  • the axial border of the second leg 126 of the cooling passageway 110A is spaced a circumferential distance D 2 from an axial border of the third leg 130 of an adjacent cooling passageway 1 IOC.
  • the distance D 2 is zero in one example, in which case the axial borders of the cooling passageways 110A, HOC are axially aligned.
  • Figure 7 illustrates Z-shaped passageways 134.
  • the Z-shaped cooling passageways 134 include an inlet 136 and a plurality of sloped portions 138, 140, and 142.
  • the sloped portions like the above-discussed sloped portions, are inclined to extend non-parallel to the radially outer surface 106.
  • the sloped portions 138, 140, 142 direct a secondary cooling airflow S in both an axial direction A and a radial direction A toward an exit 137.
  • the Z-shaped passageways further include a first leg 144 extending circumferentially between the first and second sloped portions 138, 140, and a second leg 146 extending circumferentially between the second sloped portion 140 and the third sloped portion 142. As mentioned above relative to the embodiment of Figures 4-5, the first and second legs 144, 146 may also be sloped.
  • FIG. 8 illustrates another cooling passageway 148.
  • the cooling passageway 148 is an M- shaped cooling passageway.
  • the cooling passageway 148 includes a first inlet 150, a second inlet 152, and a common exit 154.
  • a flow of fluid enters the first inlet 150, it is directed along a first sloped portion 156, turned circumferentially at a first leg 158, and directed along a second sloped portion 160.
  • Another, separate flow similarly travels from the first inlet 152, where it converges with flow from the first inlet 150 at a third, common sloped portion 162, which finally directs the converging flows from the first and second inlets 150, 152 to a common exit 154 in the radially inner surface 108.
  • the main body portion 86 may include one or more different cooling passageways.
  • one example layout is illustrated in Figure 9.
  • the main body portion 86 includes a plurality of Z-shaped passageways 134 along both the fore and aft edges 82, 84 thereof.
  • the circumferential edges 102, 104 in this example include M-shaped cooling passageways 148.
  • the Z-shaped and M-shaped passageways 134, 148 define a perimeter adjacent the outer edges of the main body portion 86.
  • a plurality of three-dimensional spiral passageways 110 are provided within the perimeter of Z-shaped and M-shaped passageways 134, 148.
  • the illustrated arrangement is particularly beneficial because it provides the inlets to each of the passageways 110, 134, 148 at a point that is spaced inward from one of the edges 82, 84, 102, 104. This inward spacing of the inlets allows for a clearance between the inlets and the adjacent engine and BOAS structures (e.g., such as the attachment hooks 88, 90, 92, 94, 96, 98, 100 in Figure 3) which may interfere with the secondary cooling flow S.
  • BOAS structures e.g., such as the attachment hooks 88, 90, 92, 94, 96, 98, 100 in Figure 3
  • FIG. 10A illustrates a divergent cooling passageway 164 which has an inlet 166 and a divider wall 168 downstream therefrom which separates a flow of fluid into two parallel flows moving along parallel sloped passageways 170, 172.
  • the passageways 170, 172 then merge and exit out the exit 174.
  • FIG 10B illustrates a U-shaped cooling passageway 176.
  • the cooling passageway 176 includes an inlet 178, a first sloped portion 180, a first leg 182 and a second sloped portion 184 which leads to an exit 186.
  • Figure IOC shows an angled cooling passageway 181.
  • the cooling passageway 181 includes an inlet 183, a sloped portion 185, and an angled portion 187 leading to an exit 189.
  • the angled portion 187 extends in a direction inclined relative to the radially outer surface 106, and relative to the direction the sloped portion 185 extends.
  • additional passageways come within the scope of this disclosure.
  • the cooling passageways described herein can be formed using any known technique.
  • One known technique includes additive manufacturing.
  • Another known technique includes investment casting.
  • a wax pattern of the BOAS segment 80 is formed.
  • a casting article e.g., a core insert
  • a wax pattern of the BOAS segment 80 is formed.
  • FIG. 12 An example casting article 192 is illustrated in Figure 12.
  • the casting article 192 is a dimensional negative of the cooling passageway 110.
  • Figure 12 is labeled with numbers corresponding to the respective portions of the cooling passageway 110, appended with a "C.”
  • C the portions of the cooling passageway 110 described above will not be repeated herein relative to the casting article 192.
  • the casting article 192 in this example is a refractory metal core (RMC) insert.
  • the RMC core may be additively manufactured.
  • the article may be a ceramic insert.
  • the casting article is provided in the wax pattern and remains part of the wax pattern until the component is cast. As is known in the art, the casting is completed, and the main body portion 86 is provided with the intended passageway.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

Un mode de réalisation, donné à titre d'exemple, de la présente invention concerne un moteur à turbine à gaz comprenant un composant présentant un corps. Ce corps comprend un passage de refroidissement sinueux, qui crée un chemin d'écoulement s'étendant entre une entrée dans une première surface du corps et une sortie dans une seconde surface du corps.
PCT/US2014/056302 2013-09-18 2014-09-18 Passage de refroidissement sinueux pour composant de moteur WO2015042262A1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
US15/023,309 US10196931B2 (en) 2013-09-18 2014-09-18 Tortuous cooling passageway for engine component
EP14845824.3A EP3047113B1 (fr) 2013-09-18 2014-09-18 Passage de refroidissement sinueux pour composant de moteur

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201361879205P 2013-09-18 2013-09-18
US61/879,205 2013-09-18

Publications (1)

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WO2015042262A1 true WO2015042262A1 (fr) 2015-03-26

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3545172A4 (fr) * 2016-11-23 2020-06-24 General Electric Company Structure de refroidissement pour élément de turbine

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10989070B2 (en) * 2018-05-31 2021-04-27 General Electric Company Shroud for gas turbine engine
CN117869016B (zh) * 2024-03-12 2024-05-17 中国航发四川燃气涡轮研究院 一种降低涡轮外环导热的冷却单元及其分析方法

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4280792A (en) * 1979-02-09 1981-07-28 Avco Corporation Air-cooled turbine rotor shroud with restraints
US20030235494A1 (en) * 2002-06-19 2003-12-25 Draper Samuel David Linked, manufacturable, non-plugging microcircuits
US20060140753A1 (en) * 2004-12-29 2006-06-29 United Technologies Corporation Blade outer seal with micro axial flow cooling system
US20070248462A1 (en) * 2005-09-30 2007-10-25 United Technologies Corporation Multiple cooling schemes for turbine blade outer air seal
US20120076645A1 (en) * 2010-09-29 2012-03-29 Rolls-Royce Plc Endwall component for a turbine stage of a gas turbine engine

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6247896B1 (en) * 1999-06-23 2001-06-19 United Technologies Corporation Method and apparatus for cooling an airfoil
US7137776B2 (en) * 2002-06-19 2006-11-21 United Technologies Corporation Film cooling for microcircuits
US6905302B2 (en) * 2003-09-17 2005-06-14 General Electric Company Network cooled coated wall
JP4634528B1 (ja) * 2010-01-26 2011-02-23 三菱重工業株式会社 分割環冷却構造およびガスタービン

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4280792A (en) * 1979-02-09 1981-07-28 Avco Corporation Air-cooled turbine rotor shroud with restraints
US20030235494A1 (en) * 2002-06-19 2003-12-25 Draper Samuel David Linked, manufacturable, non-plugging microcircuits
US20060140753A1 (en) * 2004-12-29 2006-06-29 United Technologies Corporation Blade outer seal with micro axial flow cooling system
US20070248462A1 (en) * 2005-09-30 2007-10-25 United Technologies Corporation Multiple cooling schemes for turbine blade outer air seal
US20120076645A1 (en) * 2010-09-29 2012-03-29 Rolls-Royce Plc Endwall component for a turbine stage of a gas turbine engine

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See also references of EP3047113A4 *

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3545172A4 (fr) * 2016-11-23 2020-06-24 General Electric Company Structure de refroidissement pour élément de turbine

Also Published As

Publication number Publication date
EP3047113A4 (fr) 2017-07-12
US10196931B2 (en) 2019-02-05
EP3047113A1 (fr) 2016-07-27
EP3047113B1 (fr) 2024-01-10
US20160237852A1 (en) 2016-08-18

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